# Airfoil CL & CD Start at normal magnitude then skyrocket.

I'm working with a few airfoils, NACA 23012, NACA 63015A, NACA 4412, RC4-10, and Boeing VR-7, analyzing the lift, drag and moment of each.

Currently working on the NACA 63015A. Using velocities(mph) with their respective lift/drag/moment coefficients are below.

• 25mph {-0.900/0.0257/0.0225}
• 50 {-3.7196/0.881/0.0856}
• 75 {-8.507/0.182/0.1837}
• 100 {-15.368/0.30669/0.3003}
• 125 {-24.249/0.4598/0.444}
• 140 {-21.468/1.4205/0.8104}
• 150 N/A

After 25mph both lift and drag are too high. I recently changed to reference values to get the 25mph to be in its normal range.

Airfoils' length are 8.9in = .22606m

Here are the details for the 2D case:

Solver:

Pressure-Based

Velocity formulation -> Absolute

Model:

Viscous Model -> Spalart-Allmaras (1eqn)

Vorticity-Based

Fluid:

Air

Density = 1.225 (kg/m^3)

Viscosity = 1.802e-5 (Kg/m-s)

Reference Values changes:

Area = 0.22606

Density = 1.225

Depth = 1

enthalpy = 0

Length = 0.22606

Pressure = 0

Temperature= 288.16

Velocity = varies

Viscosity = 1.802e-5

Ratio of specific heat = 1.4

Yplus for heat tran. coef. = 300

Turbulence

Specification Method -> Turbulent Viscosity Ratio

Ratio = 10

Outlet pressure = 0

Solution Method

Coupled

Gradient -> Least Squares Cell Based

Pressure->Second Order

Momentum -> Second Order Upwind

Modified Turbulent Viscosity -> Second Order Upwind

Pseudo Transient ->off

Solution Controls:

Flow Courant Number = 200 (recently reduced it to 1 after reading up on it.)

Momentum = .5

Pressure = .5

Relaxation factors

Density, Body Force, Turbulent viscosity = 1

Body Forces = .8

Report definitions of drag

x=cos(angle)

y=sin(angle)

Report definitions of lift

x=-sin(angle)

y=cos(angle)

Report definitions of moment

Center

x=1/4 of span

y= location of x in y direction.

ONE LAST THING

The contour graphics are showing the boundaries of the elements instead of a filled in contoured graph.

More pictures are below that may be relevant

Tagged:

• Hello,

Firstly, there is an abrupt jump in your mesh just beyond the trailing edge.

Secondly, please check if Reference values. Force coefficients use reference density, velocity, and area. Please check these values (especially when you are comparing with literature). Also, check if your directions are specified correctly.

Regarding the contour plot, please deselect the 'Fluid' region when you are plotting the contour. Deselect everything when you are plotting the contour plot.

Here is an example from our Ansys Innovation Courses which shows how to set-up an airfoil problem.

I hope this helps.

Thanks.

Karthik

• Hi Kartik,

The x spacing in the mesh at the trailing edge changes but the y stays the same. I checked the mesh and it didn't give me any errors.

The changes in the reference values was the problem with the 63015a airfoil at 10 degrees. I'll try this on the other AoAs and airfoils and will come back to tell you if it worked.

Thank you!

• Sure thing! Please let me know here is this worked for you.

Thank you.

Karthik